As technology progresses in the development of space, the need for orbital space structures continues to grow. Such structures will form the basic building blocks for military antenna structure and the NASA space station effort. With the high cost of transporting and assembling these structures, more and more emphasis is being placed on creating structures which can be transported in a minimum number of launches and assembled with a minimum amount of human effort.
Currently, there are two types of deployable space structures. The first type is structures whose components are fabricated and assembled on earth and subsequently deployed in space, while the second type is composite structures that are assembled in orbit with raw materials transported to the site by a launch vehicle, e.g., the space shuttle. The latter approach requires that processing equipment be lofted into orbit. The processing equipment for advanced composite materials (e.g, pultruders, filament winders, etc.) is typically weight intensive. The benefit of this approach, however, is that the raw materials may be efficiently stowed or packed in the launch vehicle without the concern of costly, elaborate and space consuming supports that terrestrially assembled structures require for transport.
Structures fabricated on earth employ two subsets of deployment systems. The first relies on mechanical strain energy stored in the system. These structures deploy by a release of boundary conditions or end constraints that allow the structure to deploy and seek a minimum potential energy configuration. The advantage of this approach is that joint and hinge type movement is minimized since most of the deployment motion occurs within the structure itself. The major drawback is that the energy storage capacity of the structure is limited, by the modulus and strength of the structural material. Furthermore, when in the stored configuration, since the internal loads are high, stress relaxation or creep may occur. This results in lost energy potential and the desired deployed configuration may not be achieved. It should also be noted that these systems are usually assisted by mechanical actuators such as power screws, friction rollers or planetary gears. Two such examples are the development of the Astromast by Astro Research Corporation, and the CTM biconvex mast developed by ESTEC, in the Netherlands.
The second subset of deployment schemes includes telescoping booms, folding plate designs, scissor mechanisms, articulated arms, and umbrella type devices. This is the most common type of design. The primary disadvantage of these designs is that they require a large number of hinges, latches, sockets, and joints. Since load demands on hinges and joints are substantial, they are areas of lowest reliability and failure during deployment.
There have been a number of recent designs that fall into this category. M. Aguirre-Martinez of ESTEC describes a thin walled carbon fiber telescoping design that requires a ball screw and motor drive to deploy. Flat panel hinged designs have been developed by D.A. Corbett and W. J. Dean of TRW Systems Group that utilize graphite/epoxy sytems. A composite tetratruss cell has been developed by M.J. Robinson at McDonnell Douglas using graphite/epoxy systems, which require the use of external actuators. Conventional scissor designs have been investigated for use in heat radiator applications by Roy L. Cox et al. of Vought Corporation. In addition, an umbrella deployment device using advanced composite ribs and wire mesh has been developed by M. Sullivan and B. McIntosh of Harris Corporation that requires a motor to open or drive the system into its deployed shape. All of these structures require external actuators that are weight and power consuming.
In common with the previously mentioned designs, new structural designs must face a multitude of requirements and constraints. The structures to be transported to the orbital site must reside in the shuttle bay, and must have a high packing density and be compact in the stowed configuration. Furthermore, while in this configuration, they must survive lift-off loads and accelerations. At the orbital site, the structure must be deployed reliably and without problems occurring due to mechanical actuator failure or human error or fatigue. Once reliably deployed at the site, the structure is required to perform its specific function. The deployed configurations are usually composed of a primary structure, such as a space truss that serves as a mechanical platform for a secondary structural component, such as an antenna, that might be a wire mesh or solar array panel. In this deployed state, there are many system considerations such as antenna surface or search control, vibrational control that requires specific damping characteristics and the consideration of long term effects of radiation or particle impact. Furthermore, extreme temperatures are experienced during orbit that induce thermal strains causing loads or defocussing deflections to occur. Therefore, these structures must have either near zero, or well characterized coefficients of thermal expansion. If residual internal loads exist in the structure after deployment, then material responses such as creep must also be considered.
The lifetime of these structures is expected to be in the order of 7 to 10 years before major modifications or maintenance must be performed. Compatible configurations and component commonality must also be considered for future structural additions. Finally, since deployment reliability and configuration accuracy is of paramount importance, structures that minimize the total number of components and rely on simple means for deployment are attractive.